In a gas turbine aircraft engine air enters at the engine inlet and flows from there into the compressor. Compressed air flows to the combustor where it is mixed with injected fuel and the fuel-air mixture is ignited. The hot combustion gases flow through the turbine. The turbine extracts energy from the hot gases, converting it to power to drive the compressor and any mechanical load connected to the drive. These hot gases produce temperature differentials that cause plastic deformation in the components exposed thereto.
The turbine consists of a plurality of stages. Each stage is comprised of a rotating multi-bladed rotor and a nonrotating multi-vane stator. The blades of the rotor are circumferentially distributed on a disk for rotation therewith about the disk axis. The stator is formed by a plurality of nozzle segments which are butted end to end to form a complete ring. Each nozzle segment comprises a plurality of generally radially disposed vanes supported between inner and outer platforms. Each vane and blade is of airfoil section.
The abutting outer platforms of the nozzle segments and the abutting outer platforms of the rotor blades collectively define a radially inwardly facing wall of an annular gas flow passageway through the engine, while the abutting inner platforms of the nozzle segments and the abutting inner platforms of the rotor blades collectively define a radially outwardly facing wall of the annular gas flow passageway. The airfoils of the rotor blades and nozzle guide vanes extend radially into the passageway to interact aerodynamically with the gas flow therethrough.
During operation of the gas turbine engine, it is desirable to minimize thermally induced plastic deformation of the outer casing. This can be accomplished by isolating the outer casing from the heat produced by the hot gases flowing through the turbine.
One technique for thermally isolating a portion of the outer casing of a turbine which surrounds a stator stage is disclosed in U.S. Pat. No. 3,644,057 to Steinbarger. According to this teaching, a heat shield encircles the outer shroud ring. The heat shield is inserted in a pair of grooves formed between the casing and outer shroud ring, which grooves constrain the ends of the heat shield against radial and axial expansion. This arrangement has the disadvantages that the heat shield will undergo plastic deformation when heated and is difficult to install in the turbine.
In U.S. Pat. No. 3,730,640 to Rice et al., a ring having heat shielding properties has a portion arranged between the outer shroud of a row of guide vanes and the outer casing. At one end the ring has a radial flange bolted to one flange on the casing and at the other end the ring has a cylindrical flange, the radially outwardly facing surface of which abuts another flange on the casing. This arrangement is disadvantageous because the ring is constrained against both axial and radial displacement by the casing flanges at two axial positions, giving rise to plastic deformation during heating.